1. Field of the Invention
The present invention relates to a gas turbine tail tube seal which seals a space between a tail tube of a combustor and a combustion gas flow channel of a turbine, and a gas turbine using the same.
2. Description of Related Art
An ordinary gas turbine is provided with a compressor and a combustor (not shown in the drawings). According to such a gas turbine, compressed air which is compressed in the compressor is supplied to the combustor and mixed with a fuel which is supplied separately so as to be combusted. The combustion gas which is generated by such a combustion is supplied to the turbine so as to generate a rotational driving force in the turbine.
A structure of an ordinary combustor is shown in FIG. 4. In the drawing, reference numeral 1 indicates a combustor, and the combustor 1 is fixed inside a wheel chamber 2. Reference numeral 3 indicates a pilot fuel nozzle to which a pilot fuel for ignition is supplied. Reference numeral 4 indicates a plurality (for example, eight) of main fuel nozzles which are disposed circumferentially around the pilot fuel nozzle 3. Reference numeral 5 indicates an inner tube. Reference numeral 6 indicates a tail tube which introduces a high temperature combustion gas F into a tail tube outlet 6a. Reference numeral 7 indicates a bypass tube. Reference numeral 8 indicates a bypass valve which forms a flow channel and opens in order to introduce air inside the wheel chamber 2 into the combustor 1 when air for combustion is insufficient due to load fluctuations. Reference numeral 9 is a seal section which is disposed in an end section of the tail tube outlet 6a and seals connecting section of a combustion gas flow channel 10 (gas path) near the turbine. A plurality of such combustors 1 are disposed around a rotor (not shown in the drawing) in the wheel chamber 2. The high temperature combustion gas F which is ejected from these combustors expands in the combustion gas flow channel 10 so as to rotate the rotor.
In the combustor 1 having the above-mentioned structure, the fuel from the main fuel nozzle 4 is mixed with air which is taken in from the surroundings and ignited by a flame of the pilot fuel which is ejected from the pilot fuel nozzle 3 so as to be combusted to become the combustion gas F. The combustion gas F passes through the inner cylinder 5 and the tail tube 6 so as to be supplied to the gas path 10 via the tail tube outlet 6a. The wall of the inner tube 5 and the tail tube 6 of the combustor 1 is always in contact with the high temperature combustion gas F. Therefore, cooling air flows in a cooling path provided in the wall. Also, the tail tube outlet 6a is connected to an inlet of the combustion gas flow channel 10 via the seal section 9; thus, the seal section 9 is cooled by the air.
FIG. 5 is a magnified cross-sectional view of section A in FIG. 4 and shows the structure of a conventional tail tube seal. As shown in the drawing, a flange 6a1 is formed around the tail tube outlet 6a. Also, the wall of the tail tube 6 is cooled by the cooling air which flows in the flow channel (not shown in the drawing) which is formed in the wall. Also, a groove 6b which follows the cooling air to flow is formed around the tail tube outlet 6a; thus, the cooling air flows therein so as to cool the wall of the tail tube 6.
The tail tube outlet 6a is connected to the combustion gas flow channel 10 via a tail tube seal 11. A groove 11a having a U-shaped cross section is formed in an end section of the tail tube seal 11. The flange 6a1of the tail tube outlet 6a is fit into the groove 11a. A groove having a rectangular cross section is formed in another end section of the tail tube seal 11. A flange section 13a of an outer shroud 13 and a flange section 14a of an inner shroud 14 of a first row stationary blade 12 in the combustion gas flow channel 10 fit into the groove 11b so as to seal the connection section.
A plurality of cooling holes 11c are formed in the tail tube seal 11 for allowing cooling air c to flow in order to cool the tail tube seal 11 and a film of the outer shroud 13 and the inner shroud 14 in the first row stationary blade 12. These cooling holes 11c are formed in an inclined manner over the outer surface and the inner surface of the tail tube seal 11. The wall of the tail tube seal 11 is cooled by allowing the cooling air c, which is highly compressed in the wheel chamber 2 to flow, to these cooling holes 11c. Furthermore, the combustion gas F which is ejected from each cooling hole 11c to the combustion gas flow channel 10, covers the inner surface of the outer shroud 13 and the outer surface of the inner shroud 14 to cool the film.
However, in the conventional tail tube seal 11, there have been the following problems.
That is, there was a concern that the film cooling operation is not effective in the conventional tail tube seal 11. Also, there was a case in which an upstream section of the outer shroud 13 and the inner shroud 14 was damaged by heat, and there was a case in which the fitting section, which fits the tail tube seal 11 to the outer shroud 13 and the inner shroud 14 became worn due to a repetitive thermal expansion. These problems are believed to occur due to the following reasons (1) to (3).
(1) The dimension of the space between the downstream end of the tail tube seal 11 to the upstream end of the outer shroud 13 and the inner shroud 14 must be large when taking the thermal expansion of each component into account. Therefore, the positions of the upstream end of the outer shroud 13 and the upstream end of the inner shroud 14 become far from the position of each cooling hole 11e. 
(2) The cooling air c is ejected form each cooling hole 11c so as to enter the combustion gas flow channel 10 in an inclined manner. Therefore, the flow of the cooling air c for cooling the film cannot be formed well along the inner surface of the outer shroud 13 and the outer surface of the inner shroud 14.
(3) The cooling air c collides with an edge of the upstream end of the outer shroud 13 and the inner shroud 14, and a vortex flow occurs when the width of the combustion gas flow channel in the tail tube seal 11 becomes larger than the width of the combustion gas flow channel made between the outer shroud 13 and the inner shroud 14 due to a reason such as thermal expansion.
The present invention was made in consideration of the above-mentioned problems. The objects of the present invention are to provide a gas turbine tail tube seal which can prevent an outer shroud and an inner shroud in a first row stationary blade from being damaged by heat and from being worn, and to provide a gas turbine having the above-mentioned gas turbine tail tube seal.